Nacelle ventilation manifold

ABSTRACT

A gas turbine engine comprises a core engine housing. A nacelle is positioned radially outwardly of the core engine housing. An outer bypass housing is positioned outwardly of the nacelle. There is at least one accessory to be cooled positioned in a chamber radially between the core engine housing and the nacelle. A manifold delivers cooling air into the chamber, and extends ng circumferentially about a central axis of the core engine. The nacelle has an asymmetric flow cross-section across a circumferential extent.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Patent ApplicationNo. 61/939,296, filed Feb. 13, 2014.

BACKGROUND OF THE INVENTION

This application relates to a ventilation manifold having a variablecross-sectional flow area across a circumference of an associatedengine.

Gas turbine engines are known and, typically, include a fan deliveringair into a bypass duct. The bypass duct is defined outwardly of a corehousing. The core housing has an upstream end, typically, known as anacelle. Cooling air may be supplied within the core housing and acrossvarious accessories that are positioned within the core housing. As anexample, various fluid components must be cooled. Also, an accessorygearbox may be positioned within the housing and must be cooled.

An inner core housing houses a compressor, a combustor and a turbinesection. The amount of air delivered into the bypass duct provides apropulsion flow in addition to an exhaust power downstream of the coreengine and which powers an associated aircraft.

The cooling air moving inwardly of the nacelle may be tapped from thebypass flow and, thus, the volume tapped reduces the efficiency oramount of propulsion provided for a given amount of fuel being burned.It is, of course, desirable to increase the efficiency.

The cooling load is not uniform across a circumference of the interiorof the nacelle. Moreover, the volume of air distributed into the nacelleat locations closer to an inlet will be greater.

In known manifolds, the flow area has typically been uniform across thecircumference of the engine.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine comprises a core enginehousing. A nacelle is positioned radially outwardly of the core enginehousing. An outer bypass housing is positioned outwardly of the nacelle.There is at least one accessory to be cooled positioned in a chamberradially between the core engine housing and the nacelle. A manifolddelivers cooling air into the chamber, and extends ng circumferentiallyabout a central axis of the core engine. The nacelle has an asymmetricflow cross-section across a circumferential extent.

In another embodiment according to the previous embodiment, a pluralityof flow distribution tubes extend from the manifold and into thechamber.

In another embodiment according to any of the previous embodiments, theplurality of flow distribution tubes includes an exit opening at an endof the flow distribution tubes and a plurality of openings formed in asidewall.

In another embodiment according to any of the previous embodiments, themanifold is defined by having a radially outer sidewall and an innerwall, with a rear wall and a radially inner wall defined by a separatehousing member.

In another embodiment according to any of the previous embodiments, aninlet tube for the manifold captures air from a bypass duct definedbetween the outer bypass housing and the nacelle.

In another embodiment according to any of the previous embodiments, themanifold has a greater cross-sectional area associated with acircumferential location where the inlet tube enters the manifold, andhas a lesser cross-sectional area at locations spaced circumferentiallyfrom the area of greater flow cross-sectional area.

In another embodiment according to any of the previous embodiments, themanifold delivers cooling air over at least two of the accessories. Afirst of the accessories is associated with the area of greatercross-sectional area, and includes an accessory gearbox. A secondaccessory is associated with the area of lesser cross-sectional area.The second accessory is a fluid component.

In another embodiment according to any of the previous embodiments, avalve is controlled to modulate the volume of air delivered as coolingair into the chamber from the manifold.

In another embodiment according to any of the previous embodiments, thevalve is open to allow greater airflow to cool the at least onecomponent during high pressure operation of an associated gas turbineengine and is modulated to reduce the airflow at lesser power portionsof the gas turbine engine.

In another embodiment according to any of the previous embodiments, thevalve is located on an inlet tube leading into the manifold.

In another embodiment according to any of the previous embodiments, themanifold is defined by having a radially outer sidewall and an innerwall. A rear wall and a radially inner wall are defined by a separatehousing member.

In another embodiment according to any of the previous embodiments, aninlet tube for the manifold captures air from a bypass duct definedbetween the outer bypass housing and the nacelle.

In another embodiment according to any of the previous embodiments, themanifold has a greater cross-sectional area associated with acircumferential location where the inlet tube enters the manifold, andhas a lesser cross-sectional area at locations spaced circumferentiallyfrom the area of greater flow cross-sectional area.

In another embodiment according to any of the previous embodiments, themanifold delivers cooling air over at least two of the accessories. Afirst of the accessories is associated with the area of greatercross-sectional area, and includes ng an accessory gearbox. A secondaccessory is associated with the area of lesser cross-sectional area.The second accessory is a fluid component.

In another embodiment according to any of the previous embodiments, avalve is controlled to modulate the volume of air delivered as coolingair into the chamber from the manifold.

In another embodiment according to any of the previous embodiments, thevalve is open to allow greater airflow to cool the at least onecomponent during high pressure operation of an associated gas turbineengine and is modulated to reduce the airflow at lesser power portionsof the gas turbine engine.

In another embodiment according to any of the previous embodiments, thevalve is located on an inlet tube leading into the manifold.

In another embodiment according to any of the previous embodiments, acompressor and a turbine are mounted within the core engine housing.

In another featured embodiment, a gas turbine engine comprises a coreengine housing. A nacelle is positioned radially outwardly of the coreengine housing, and an outer bypass housing is positioned outwardly ofthe nacelle. There is at least one accessory to be cooled positioned ina chamber radially between the core engine housing and the nacelle. Amanifold delivers cooling air into the chamber. A valve is controlled tomodulate the volume of air delivered as cooling air into the chamberfrom the manifold.

In another embodiment according to the previous embodiment, the valve isopen to allow greater airflow to cool the at least one component duringhigh pressure operation of an associated gas turbine engine and ismodulated to reduce the airflow at lesser power portions of the gasturbine engine.

In another embodiment according to any of the previous embodiments, thevalve is located on an inlet tube leading into the manifold.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows the location of a nacelle ventilation manifold.

FIG. 3 shows a detail of the manifold.

FIG. 4 shows a portion of a manifold.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption ('TSFC')”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows an interior chamber 100 inward of a nacelle housing 108.This may be roughly the chamber identified by 11 in FIG. 1, and betweennacelle housing 12 and core housing 13. An inlet pipe 110 has an opening14 that captures air from the bypass duct (B) and delivers it into abody of a manifold 104. The air is then distributed outwardly through aplurality of flow distribution tubes 106. The air delivered outwardly ofthe flow distribution tubes 106 cool components within the nacellehousing, including an accessory gearbox 102. As known, the accessorygearbox includes gears which are driven by a drive shaft of a gasturbine engine such as engine 20 of FIG. 1. The accessories may includean oil pump, a fuel pump, etc. The accessory gearbox 102 carries a highcooling load as it generates a good deal of heat.

The air leaving the tubes 106 also cools fluid components such as shownby tube 89 in FIG. 2. The cooling load may be greater near opening 14 ofthe engine, such as the lowermost portion in the FIG. 2 orientation.

As shown in FIG. 3, the inlet pipe 110 and associated pipe 99 enters alower area 114 of the manifold 104, which will be aligned with thegearbox 102. The lower portion 114 will then flow in both clockwise 111and counterclockwise 113 directions to more remote areas of the interiorof the manifold 104. As can be appreciated in FIG. 3, the disclosedmanifold 104 may include an outer sidewall 199 and a forward end wall101.

Returning to FIG. 2, portions 132 and 131 of an internal housing mayclose off the other two walls to define an interior flow path 200 withinthe manifold 104.

At upper locations 116, one can appreciate that a cross-sectional areaof the flow path 200 will be smaller than it will be at the lowerportion 114. The manifold is asymmetric and in flow cross-section acrossa circumferential extent about center line A.

This asymmetric sizing provides two functions. First, the greater sizemay be provided adjacent the area of greatest heat load, as an example,the accessory gearbox 102. In addition, the air at the lower portion 114will exit through tubes 106 and also flow in directions 111 and 113towards more remote tubes 106. As such, the volume of air reaching theupper areas 116 will be smaller than the volume of air beginning at thelocation 114. The asymmetric sizing facilitates the flow of the air inthose directions.

A valve 300 allows modulation of the nacelle cooling flow to provideperformance benefits. As an example, a higher volume of airflow may beallowed by opening the valve 300 during high power flight portions, suchas takeoff. On the other hand, lower power portions, such as cruise, mayhave the volume of airflow reduced dramatically and even stopped. Thiswill result in more efficient use of the bypass airflow. A control 302is shown schematically for controlling the valve 301.

FIG. 4 shows details of the flow distribution tubes 106. There is an endopening 114 and a plurality of openings 112 in a sidewall 113.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a core engine housing; a nacellepositioned radially outwardly of said core engine housing, and an outerbypass housing positioned outwardly of said nacelle; there being atleast one accessory to be cooled positioned in a chamber radiallybetween said core engine housing and said nacelle; and a manifold fordelivering cooling air into said chamber, said manifold extendingcircumferentially about a central axis of said core engine, and saidnacelle having an asymmetric flow cross-section across a circumferentialextent.
 2. The gas turbine engine as set forth in claim 1, wherein aplurality of flow distribution tubes extend from said manifold and intosaid chamber.
 3. The gas turbine engine as set forth in claim 2, whereinsaid plurality of flow distribution tubes include an exit opening at anend of said flow distribution tubes and a plurality of openings formedin a sidewall.
 4. The gas turbine engine as set forth in claim 2,wherein said manifold is defined by having a radially outer sidewall andan inner wall, with a rear wall and a radially inner wall defined by aseparate housing member.
 5. The gas turbine engine as set forth in claim2, wherein an inlet tube for said manifold capturing air from a bypassduct defined between said outer bypass housing and said nacelle.
 6. Thegas turbine engine as set forth in claim 6, wherein said manifold has agreater cross-sectional area associated with a circumferential locationwhere said inlet tube enters said manifold, and having a lessercross-sectional area at locations spaced circumferentially from saidarea of greater flow cross-sectional area.
 7. The gas turbine engine asset forth in claim 6, wherein said manifold delivering cooling air overat least two of said accessories, with a first of said accessoriesassociated with said area of greater cross-sectional area, and includingan accessory gearbox, and a second accessory associated with said areaof lesser cross-sectional area, said second accessory being a fluidcomponent.
 8. The gas turbine engine as set forth in claim 2, wherein avalve is controlled to modulate the volume of air delivered as coolingair into the chamber from the manifold.
 9. The gas turbine engine as setforth in claim 8, wherein said valve is open to allow greater airflow tocool said at least one component during high pressure operation of anassociated gas turbine engine and is modulated to reduce the airflow atlesser power portions of the gas turbine engine.
 10. The gas turbineengine as set forth in claim 9, wherein said valve is located on aninlet tube leading into said manifold.
 11. The gas turbine engine as setforth in claim 1, wherein said manifold is defined by having a radiallyouter sidewall and an inner wall, with a rear wall and a radially innerwall defined by a separate housing member.
 12. The gas turbine engine asset forth in claim 1, wherein an inlet tube for said manifold capturingair from a bypass duct defined between said outer bypass housing andsaid nacelle.
 13. The gas turbine engine as set forth in claim 1,wherein said manifold has a greater cross-sectional area associated witha circumferential location where said inlet tube enters said manifold,and having a lesser cross-sectional area at locations spacedcircumferentially from said area of greater flow cross-sectional area.14. The gas turbine engine as set forth in claim 1, wherein saidmanifold delivering cooling air over at least two of said accessories,with a first of said accessories associated with said area of greatercross-sectional area, and including an accessory gearbox, and a secondaccessory associated with said area of lesser cross-sectional area, saidsecond accessory being a fluid component.
 15. The gas turbine engine asset forth in claim 1, wherein a valve is controlled to modulate thevolume of air delivered as cooling air into the chamber from themanifold.
 16. The gas turbine engine as set forth in claim 15, whereinsaid valve is open to allow greater airflow to cool said at least onecomponent during high pressure operation of an associated gas turbineengine and is modulated to reduce the airflow at lesser power portionsof the gas turbine engine.
 17. The gas turbine engine as set forth inclaim 15, wherein said valve is located on an inlet tube leading intosaid manifold.
 18. The gas turbine engine as set forth in claim 1,wherein a compressor and a turbine are mounted within said core enginehousing.
 19. A gas turbine engine comprising: a core engine housing; anacelle positioned radially outwardly of said core engine housing, andan outer bypass housing positioned outwardly of said nacelle; therebeing at least one accessory to be cooled positioned in a chamberradially between said core engine housing and said nacelle; and amanifold for delivering cooling air into said chamber, a valve iscontrolled to modulate the volume of air delivered as cooling air intothe chamber from the manifold.
 20. The gas turbine engine as set forthin claim 19, wherein said valve is open to allow greater airflow to coolsaid at least one component during high pressure operation of anassociated gas turbine engine and is modulated to reduce the airflow atlesser power portions of the gas turbine engine.
 21. The gas turbineengine as set forth in claim 19, wherein said valve is located on aninlet tube leading into said manifold.